Aircraft automatic pilot



Jan. 7, 1964 G. F. JUDE ETAL AIRCRAFT AUTOMATIC PILOT 6 Sheets-Sheet lOriginal Filed May 14, 1959v ATTORNEY Jan. 7, 1964 G. F. JUDE ETAL315899 AIRCRAFT AUTOMATIC PILOT Original Filed May 14, 1959 6Sheets-Sheet 2 Jan. 7, 1964 G. F. JUDE ETAL AIRCRAFT AUTOMATIC PILOTOriginal Filed May 14, 1959 6 Sheets-Sheet 3 Jan. 7, 1964 G. F. JUDEETAL AIRCRAFT AUTOMATIC PILOT original Filed May 14, 1959 6 Sheets-Sheet4 INVENTQRS PGE F. JUDE /PRY MILLER ATTQRNE;I

AIRCRAFT AUTOMATIC PILOT Original Filed May 14, 1959 Jan. 7, 1964 G. F.JUDE E-rAl.l

6 Sheets-Sheet 6 INVENTORS GEORGE F. JUDE BYHARRY M/LLE'? ATTORN nitedStates The presen-t invention relates generally to improvements inautomatic control systems for aircraft and more particularly to anautomatic pilot for aircraft of the character set forth in copendingparent application Serial No. 571,813 led March l5, 1956, now Patent3,007,656 for Aircraft Automatic Pilots, and the present applicationconstitutes a continuation-impart f that application. The presentapplication contains a more detailed disclosure of the automatic pilotof the referenced parent application and is directed to improvements notspecifically disclosed therein. The present application is alsoA adivision of patent application Serial No. 75,432 led December l2, 1960,now Patent No. 3,079,109, in the names of George F. Jude and HarryMiller which in turn is a division of patent application Serial No.813,097 tiled May 14, 1959 and issued September 5, 1961 ias Patent No.2,998,946 in the names of George F. Jude and Harry Miller. All of saidapplications being assigned to the present assignee.

The various objects and feature of the improved automatic pilot of thepresent invention will become apparent from the following description ofthe preferred embodiment thereof taken in connection with theaccompanying drawings, wherein,

FIGS. 1A, 1B and 1C together illustrate the present automatic pilot inblock diagram form; and

FIGS. 2 through 14 are schematic ywiring diagrams of the interlockcircuits employed for switching theautomatic pilot from one mode ofoperation to another in response to operation of manually controlledselector switches, maneuver command switches, and the like. For thereaders convenience, the general function of each of the separatelyillustrated interlock circuits are suggested by a figure title.

"Ihe following speciiication will be divided into sections in accordanceIwith the various modes of operation availyable to and selectable by thehuman pilot. Generally these modes are Off or Disengage, Yaw Damper,Autopilot Engage, Manual Maneuver, and Path Control. Submodes ofoperation under the latter are further available such as AltitudeControl, Radio Guidance (VOR/LOC), and Approach (Glide Path). For anillustration of the outward appearance of the pilots control whereby theforegoing modes of operation may be selected, reference may be made toFIG. 7 of the above-identified parent application. During thedescription of the autopilot in its various modes of operation, thefunctioning of the interlock circuits, which in some cases automaticallychanges the mode of operation of the autopilot and which also prohibitsincompatible signals from influencing pilot operation in any particularmode, will be set forth.

n the drawings, all relays are considered to be deenergized and lallrelay-controlled switches are accordingly illustrated in theirdeenergized positions. The switch contacts represented by an arrow ortriangle denote those contacts -which the switch armatures engage whentheir controlling relays are energized.

Gen eral Referring now to IFIGS. 1A, 1B and 1C, which are so ICCarranged that they may be conveniently fixed together, it will be seenthat the autopilot comprises generally three channelsa rudder channel109, an aileron channel 260, and an elevator channel 3681. Each of theserudder, aileron, and elevator channels may be broken down generally intothree sections, namely, a short period stabilization section designated161, 201 and 331, respectively, a long period stabilization sectiondesignated `182, 262 and 3t12, respectively, and a command section, thelatter being further broken down into a manual input subsection 203 forlateral command and 303 for longitudinal command, and an externalreference input subsection designated 4% for the lateral externalcommand and 491 and 401' for longitudinal external commands.

Each of the three channels terminates in a substantially identicalservomotor system 104, 204 and 3nd for controlling the movement of therudder, aileron, and elevator, respectively. These servomotor systemsare the same as those described in detail in the above-identified parentapplication and -will not be described in detail herein. Sufce it to saythat they are velocity type servos, that is, during the normal operationof the automatic pilot the servos are caused to drive in a direction andat a rate proportional to the direction and magnitude of the inputsignal suppled thereto. This is accomplished by means of a rategenerator driven by a servomotor whose output is fed back to the inputof a high gain servo amplifier in opposition to the control signalapplied to the amplifier. It will be noted that there is no positionfeedback during normal operation. The aircraft loop is closed by meansof direct measures of craft langular acceleration, i.e. the surfacedeection is controlled by the angular acceleration such deflectionproduces on the aircraft. Such arrangement provides substantiallyoptimum control of the aircraft under all flight conditions. As will bedescribed below, during the Disengage mode of operation, servomotorposition information is required and for this purpose a synchro isclutched to the motor output in this mode. The elevator servo systemincludes a trim tab control system 364 for operating the trim tab of theaircraft for automatic trim purposes as also described in the parentapplication.

As in the parent applica-tion, automatic short period stabilization ofthe aircraft about each of the crafts primary axes while under controlof the automatic pilot is provided by a pair of linear accelerometersspaced a distance apart along each axis and their combined outputs areso connected that a signal is produced which is proportional to theangular acceleration of the line joining the accelerometers about anaxis at right angles to such line. Also, as in the parent application,one of the accelerometers in the yaw and pitch channel is mounted asnear as possible to the C.G. of the aircraft and its output isolated insuch a manner from its` companion accelerometer that such output is ameasure of the lateral and vertical linear accelerations of theaircraft, respectively. The manner in which the 'accelerometers aremounted in the aircraft and electrically connected is illustrated indetail in the parent application. In the present application, referencecharacter designates the forward yaw accelerometer and 106 designatesthe C.G. lateral accelerometer, the output of the latter being combinedwith the output `of the former to provide the yaw angular accelerationsignal, and the `output taken alone constituting the lateralacceleration signal. Similarly, reference char- -acters 305 and 396designate respectively the forward pitch accelerometer and the C.G.vertical accelerometer respectively, the output of the latter beingcombined with the output of the former to provide the pitch angularacceleration signal, and the output of the latter alone supplying thevertical acceleration signal. In the roll channel, reference characters2615 and 206 designate respectively the upper and lower rollaccelerometers, the combined output of which provides a measure of theangular acceleration of the craft in roll, i.e. about the craftforeand-aft axis.

The long period displacement stabilization of the craft about itsprimary axes is provided by a vertical gyro 6 which supplies rollattitude data from output 207 and pitch attitude data from output 397.This vertical gyro may be a conventional gyro having, for example,liquid levels for maintaining the spin axis thereof slaved to gravityvertical. The gyro may also be provided with a quick erecting deviceoperable when the pilot is initially turned on. Furthermore, it may beprovided with suitable means for cutting out its erection underconditions where the effect of gravity vertical has been displaced dueto craft accelerations. A schematic representation of a vertical gyrosuitable for use herein is illustrated in the parent application.

Long period displacement stabilization is provided in the presentautomatic pilot by a directional gyroscope 107 which may be suitablyslaved to the magnetic meridian. This gyro may likewise be equipped withquick erection devices and means for cutting off its slaving duringperiods of flight, wherein accelerations may affect the magnetic datasensor to which the gyro is referenced. In the absence of externalcommands, the attitude displacement signals provided by the vertical anddirectional gyros are applied unmodified to the surfaces of all systemsas will become apparent below.

All external commands applied to the automatic pilot are supplied to theservomotor systems through command computers which function generally tosmooth the input command signal so that abrupt transients are eliminatedor in some modes as electro-mechanical time integrators. Thus, commandseffecting movement of the craft in the vertical direction are applied tothe elevator servo system through a pitch command computer 308, andcommands effecting lateral movement of the craft, i.e., turns, aresupplied to the aileron surface servomotor system through a roll commandcomputer 268. A yaw command computer 108 is operable to followup on thedirectional gyro output during turn commands.

Pitch commands to the autopilot may be applied from three sources; apilot-controlled pitch-rate command 309, displacement of which from adetent position commands a rate of change of pitch attitude whereby theresponse of the craft to such command is identical to that which wouldnormally occur by the pilots moving his control column. The secondexternal pitch command is supplied by an altitude control 310, thedetails of which are illustrated in copending application Serial No.571,788 filed March l5, 1956, for Longitudinal Axis Control System forAircraft, and assigned to the same assignee as the present application.The third source of pitch command signal is the glide slope receiver 311which supplies signals proportional to the displacement of the craftfrom the radio-defined glide path to the radio coupler 400 whichconverts the output of the receiver 311 to signals acceptable by theautomatic pilot. It should be noted that the altitude control 310 andglide slope control 311 constitute apparatus for defining a fiight pathin the vertical plane which the aircraft is caused to follow.

Turn commands for the automatic pilot may be supplied from threesources. A pilots turn knob 209 which, like the pitch knob 309, providesan output which cornmands a rate of turn proportional to the sense ofdisplacement of the turn knob from its detent position. Again, the craftresponse to such command is identical to that which would occur uponturning of the pilots wheel. The second turn command source is a pilotsmagnetic heading selector 210. Through this command the pilot may selectany desired magnetic heading and the craft will smoothly turn to andthereafter maintain the magnetic heading so selected. The third sourceof turn command is the output of a VOR/LOC radio receiver 211 which,like glide path receiver 311, provides a signal in accordance with thedisplacement of the craft from a radio defined course. The output of thelateral receiver 211 is modified, in a manner to be fully describedbelow, in radio coupler 400 the output of which is applied to theautomatic control system to cause the craft to approach and thereaftermaintain that radio beam which is tuned in on the receiver.

While there is no main Off-On switch for the autopilot, it iseffectively turned on by closure of three circuit breakers. Thesecircuit breakers are shown in FIGS. 2 and 5. An instrument circuitbreaker labeled Inst CB, an autopilot circuit breaker labeled APCB, anda yaw damper circuit breaker labeled YDCB. As illustrated, closure ofthe instrument circuit breaker serves to energize the vertical gyro andits associated quick erection system as well as to energize othercomponents of the autopilot requiring fixed phase A.C. Similarly,closure of the autopilot circuit breaker supplies energization currentfor the roll and pitch accelerometer pairs as well as other componentexcitation such as roll and pitch servo fixed phase energization, rolland pitch command computer fixed phase energization, etc. At the sametime, the autopilot circuit breaker controls an undervoltage sensorwhich, upon a loss or dropping of A.C. power below a predeterminedminimum, serves to deenergize relay K- which automatically disengagesthe roll and pitch servo systems from their respective control surfaces.Closure of the yaw damper circuit breaker supplies excitation currentfor the yaw accelerometer pair as well as other component excitationnecessary for yaw damper operation; for example, energization of therudder servo system fixed phase supply. The yaw damper circuit breakerthrough rectifier R1 supplies D.C. voltages to the autopilot and yawdamper engage interlock circuits, the operation of which will bedescribed simultaneously with the description of the autopilot in itsvarious modes of operation. Of course, the autopilot may be completelydisabled by opening all of these three circuit breakers.

Dsengage M 0de With all of the pilot circuit breakers closed but theservos disengaged, all the sensors are operative to detect movement ofthe craft about its primary axes; the yaw, roll and pitch commandcomputers 108, 208 and 308 are all placed in follow-up on the referencesignals provided by displacement gyros, i.e., the directional gyro 107and vertical gyro 6; and the surface servo systems are placed inposition follow-up on any signals applied to their input amplifiers.Thus, at any time during the disengage mode the human pilot may placethe aircraft in the engage mode of operation without abrupt transients.The disengage mode is automatically instituted by placing the servoengage switch, schematically illustrated at 9 in FIG. 2, in its Offposition. In this position selector switch arms 112, 212, and 312terminate on grunded contacts and arm 10 terminates on an open contact.Under these conditions the rudder, aileron and elevator engage solenoidsRES, AES and EES are deenergized and the clutches controlled thereby areopen. It will be noted that the direct link between the engage switchesand the engage solenoids establishes an important safety feature in thatactuation of the engage switch to its off position immediately opens thecircuit to the solenoids, thereby immediately and positively disengagingthe servos should an emergency arise.

As stated, the servos are placed in position follow-up on any inputsignal to the servo amplifier during the disengage mode. This isaccomplished through rudder, aileron, and elevator engage synchrosolenoids RSS, ASS and ESS which remain deenergized since the switch 10is on an open terminal and the main autopilot engage solenoid K-102D andyaw damper solenoid K-IZS are both deenergized, thereby opening switchesK-tl-ZB-l,

-2, -3 and K-12S-2. Deenergization of these solenoids establishes adirect connection between rudder, aileron and elevator servomotors andposition synchros 113, 213 and 313, respectively. Since, in thedisengage mode, pilot engage relay K-102D and yaw damper engage relayK-12S are both deenergized, switch K-ltiZD-l is open and relay K-102A islikewise deenergized which, in turn, causes the roll command computer29S and the pitch `command computer 39S to be placed in follow-up on theroll and pitch gyro signals through operation of switches K-102A-1 andK-102A-2- Similarly, the yaw command computer S is caused to follow-upon the directional gyro signal through the operation of yaw clamp relayK-170 which is normally energized in the engaged modes. Thus, yaw clamprelay switch K-170-1 is closed by virtue of open contact of engage`switch K-lltZB-d (FiG. 11) energizing the yaw follow-up motor.

With the command computers 168, 26 and 303 in their follow-up mode, thetachometer gradients are suitably changed so that the follow-up israpid. The roll tachometer gradient is changed by virtue of rollfollowup low tachometer gradient deenergized relay K-lSt) (FIG. 12) andits associated switch K-lt-l (FIG. 1B). Similarly, pitch follow-uptachometer gradient deenergized relay K-190 (FIG. 7), through itsassociated switch K-19ti-1, places the pitch follow-up 30S in fastfollow-up on the vertical gyro pitch signal by virtue of deenergizedengage relay K-102D. The yaw command computer is normally in fastfollow-up. However, it is placed in slow follow-up during coupling toILS beams as will be described hereinafter.

Also, in the disengage mode, disengage relay switch K-102A-3 operates toshort out any signal accumulated by the roll integrator 110.

Yaw Damper In the yaw damper mode of operation of the autopilot of thepresent invention, the aircraft is under the manual control of the pilotand the yaw damper portion of the automatic pilot only provides dampingof yaw motions of the craft. The yaw damper is especially designed todamp out any Dutch Roll tendency of the airplane during manual flight, acondition which is inherent in most modern aircraft having severelyswept wings. The yaw damper is so designed that the human pilot does notnormally known that the yaw damper is operating. In pilot commandedturns, however, he may be conscious that the yaw damper is helping himsince he need not apply any rudder pedal to coordinate his commandedturns. All he need do is to turn his aileron control wheel in thedirection it is desired to turn and in an amount dependent upon thedesired turn rate. The manner in which this is accomplished will becomeevident as the description of the yaw damper mode continues.

To institute the yaw damper mode of operation, the pilot moves his servoengage switch 9 to the Yaw Damper position (FIG. 2). It will be notedthat rectifier 8 supplies two DC. voltages. One supply is fused as byfuse F and the other is not. The fused D C. energizes relay K-Sit) whichserves as a D C. voltage sensor and normally holds its associated switchK-Stitl-l in its energized or down position. The non-fused supplyenergizes only that portion of the autopilot and yaw damper engageinterlock circuit which is necessary for yaw damper operation. If lafailure in the automatic pilot interlock channels blows fuse F, thenon-fused supply is available to operate the yaw damper. \Vith switch 9in its yaw damper position, fused D.C. is applied through the pilots andco-pilots emergency release buttons P and CP to a yaw damper engagerelay K-128 and an engage switch holding solenoid 12. The function ofthe engage switch holding solenoid 12 is to provide a mechanicallyreleasable lock on the engage switch 9 such that with the switch in theyaw damper position, should the solenoid become deenergized, the engageswitch will snap automatically through spring action to its offposition, thereby placing the autopilot in the 6 disengage mode.Energization of yaw damper relay K-128 serves to close its associatedswitch K-128-2, which energizes the rudder engage solenoid RES throughswitch arm 112 to thereby engage the rudder servomotor with the aircraftrudder. Simultaneously, the rudder synchro solenoid RSS disengages therudder synchro 113 from the servo system (FIG. 1C) so that the onlyfeedback to the servo amplifier Vis that from the servo tachometer.Clutching of the rudder engage clutch serves to close rudder microswitchRMS, thus establishing a circuit which will energize the engage switchholding solenoid 12. If for some reason the rudder engage solenoid doesnot function, or should any other malfunction occur which would cause adeclutching of the rudder servomotor from the surface, microswitch RMSopens thereby deenergizing engage switch holding solenoid 12 and causingswitch 9 to rop to its off position.

The foregoing interlock operation places the rudder servo system underthe control of yaw accelerometers and 196, as well as the rollaccelerometers 205 and 206. The two basic sensors for yaw damping arethe yaw accelerometers. Their combined outputs, proportional to yawangular acceleration, are applied to an amplifier demodulator 114, theDC. output of which is applied in two branch circuits, one containing alead circuit and filter 115 (shown in detail in the pitch channel 301)which functions as a blocking network and noise filter in theacceleration signal due to power supply fluctuations and/ or tobody-bending effects. The other branch leads the acceleration signal toa lag-lead filter 116, the lag portion of which serves to develop a ratecomponent therefrom, while the lead portion thereof serves to againsuppress noise and to compensate for any body-bending effects. Thedetails of these circuits are shown in the pitch channel. The angularacceleration and rate terms are combined and modulated in demodulator117 and the resultant output is applied to the input of the rudder servoamplifier 113 which is a very high-gain amplifier as described in theparent application. After amplification therein, the signal is used tocontrol the rudder servomotor 119 which drives the rudder at a rateproportional to the magnitude of the amplifier input by virtue oftachometer feedback from rate generator 120. Thus, angular accelerationsof the craft about the yaw axis serve to drive the control surfaces at arate proportional thereto, thereby producing a yaw acceleration of thecraft in opposition to that produced by the disturbing moment whichinitially caused the angular acceleration. The rapid response of thesurface to angular accelerations greatly increases the apparent inertiaof the airplane, thereby minimizing attitude deviations in turbulentair.

As stated, through the use of additional signals, turns produced by thehuman pilot are coordinated through yaw damper operation. Long term orsteady state coordination is provided by C.G. located accelerometer 1%.Its output is parameter controlled as at 109, rendered operative in theyaw damper mode by relay switch K-1ZS-3, and is then applied todemodulator 121, the DC. output of which is smoothed as by lag filter122 and remodulated at 123. The output of the demodulator is appliedthrough a direct connection to the rudder servo system input and also-to yaw damper integrator 124 whose output, in turn, is combined withthe angular acceleration signal and applied to the servo amplifier 118.The direct feed of later-al acceleration to the rudder servo togetherwith the yaw damper integrator output serves to provide the long term orsteady state coordination during human pilot initiated turns. Short termor transient miscoordination is substantially eliminated by driving therudder as a function of the rate of roll of Ithe aircraft. In otherwords, coordination i's provided during turn entry and turn exit or inthe event of roll producing gusts during the turn. The roll rate signalis developed as follows. The upper and lower accelerometers 205 and 266provide a signal proportional to roll angular acceleration as in theparent application, this signal being amplified and demodulated at 214,the D.C. output of which is branched to lead circuit and filter 215 andlag-lead lter 216. Lead circuit and filter 215 serves as a blockingnetwork and noise filter as in the rudder channel, while lag-lead lter216 serves to Iderive a rate component from the acceleration signal andlikewise to suppress noise as above. The roll rate component isparameter controlled at 216i and combined with the angular accelerationsignal in demodulator'217, the output of which is applied to lthe inputof rudder servo 1S through the yaw damper integrator 124 after beingparameter adjusted at 217. Through the operation of integrator 124, theroll signal applied to the rudder servo amplifier varies in accordancewith the rate of roll of the aircraft. The parameter controlling theroll rate term provides transient coordination under all ightconfigurations.

Since the aileron servo is disengaged, the roll angular accelerationsignal will be ineffective in the aileron channel. It will be noted inregard to the interlocks in the yaw damper mode that no importantprecautions need be taken prior to engaging the rudder to its servosystem inasmuch as it is only the higher derivative craft movements thatare being damped out. In other words, the yaw damper mode may beinstituted at any time, no matter what the aircrafts attitude might beat that time.

Autopzlor Engage Before discussing the autopilot mode of operation ofthe autopilot of the present invention, it should be noted that thosecircuits operative in the yaw damper mode jus-t described operate in anidentical manner in the autopilot mode so that a description thereofwill not be repeated.

In order to place the autopilot in the full autopilo mode, engageswitch' 9 is thrown to its Autopilot position, thereby moving switch arm1f) to its uppermost contact, establishing D.C. power at this point.However, before the full autopilot may be engaged, certain conditionsmust be satisfied. When all these conditions are satisfied and the pilotinterlock loop is closed, the autopilot engage relay K-102D isenergized, thereby permitting operation of the automatic pilot in allits possible modes of operation. The conditions which must be fulfilledare as follows. There must be sufiicien-t D.C. power as determined byswitch K-500-2 controlled from undervoltage relay K-SOO; the verticalgyro must be fully erected and operating normally as determined byswitch K-3il2-1 controlled from quick erection relay K-302 (FIG. 5).Also, sufficient A.C. voltage must be present in the system asdetermined by A.C. undervoltage sensor 7 and relay K-120 normallyenergized thereby. Switch K-120-1 determines if this conditon issatisfied before the pilot can be engaged. A further condition whichmust be fulfilled before pilot engage relay K-102D will be energized isthat all autopilot automatic cut-off relays will be balanced, i.e., thatthere is no condition in the autopilot which would cause -its automaticcut-off. The details of such automatic cut-oli? or safety monitor systemare disclosed in copending application Serial No. 623,592 for MonitoringDevice for Automatic Pilot Systems, filed November 2l, 1956, in thenames of H. Miller and R. H. Parker, now issued as Patent No. 2,973,927dated March 7, 1961, and `assigned to the same assignee as the presentinvention. For simplicity of illustration, these automatic cut-offrelays have been illustrated generally by a suitably labeled block 13.

Also, the autopilot cannot be engaged at bank angles greater than 35.Such condition is determined by bank detector relay K-114 (FIG. 1B).This relay is energized through the operation of the roll commandcomputer 208 as, during the disengage mode, it follows up on thevertical gyro roll signal supplied by roll synchro 218 on vertical gyro6 and synchro 219 on the roll computer output shaft. An extension of theroll computer shaft positions a switch arm 220 along an arcuate sector221 which subtends an angle of 35 on each side of zero roll'referenceposition. D.C. current is applied to the wiper 220 such that if thewiper is off the contact strip, bank detector relay K-114 is`deenergized and switch K-114-1 (FG. 2) will remain open, therebypreventing engagement of the pilot. Of course, if the bank angle is lessthan plus or minus 35, switch K-114-1 is closed permitting pilotengagement, provided the other engaged conditions are fulfilled.

A still vfurther condition which must be satisfied before the autopilotengage relay will become energized is that the manual pitch and turncontrollers must be in their no command or detent position. This isaccomplished in the engage interlock circuit for relay K-102D bynormally closed relay switches K-104B-1 and K-106-1 which, in turn, arecontrolled by turn knob detent relay K-lltM-B and pitch knob detentrelay K-106, illustrated in FG. 3. As shown in this figure, should theturn knob 14 be out of detent, switch 15 will be closed therebyenergizing relay K-lO/iB and opening engage interlock circuit switchK-IMB-l. Similarly, should the pitch knob y16 be out of detent, swt-ich17 will be closed and relay K-lto energized, thus opening K-106-1 in theengage interlock circuit (FIG. 2).

In the pilot engage interlock circuit of FIG. 2, the engage relay switchK-LlGZC-l is shown in parallel with the bank angle switch K-114-1 andthe detent switches K-lii-iB-l and K-ltt-l. This switch is closed uponenergization of the main pilot engage relay K-102D and serves to insurethat subsequent operation of the bank angle and detent relays will notdisengage the pilot.

With all of the foregoing conditions satisfied, the pilot engage relayK-102D is energized, which energization controls a number of parallelconnected engage relays K- 102A, -B, and -C through the closure ofengage relay switch K-ldZD-l. Energization of K-102B closes gangedswitches K-IOZB-l, -2 and -3 which, in turn, energize the rudder,aileron, and elevator synchro solenoids RSS, ASS and ESS to declutch thesurface position synchros 113, 213 and 313 from the servomotor outputsand simultaneously, by virtue of switch arms 112, Y212 and 312 being intheir uppermost position, energize the rudder, aileron and elevatorengage solenoids RES, AES and EES, thereby effecting direct connectionfrom the servomotors to the control surfaces. It will be noted that ifthe craft had previously been in the yaw damper mode, the rudder synchrosolenoid and rudder engage solenoids will remain energized through theoperation of engage switch K-ltlZB-S, the yaw damper relay switchK-128-2 having been open upon switching from the yaw damper to theautopilot mode. With the rudder, aileron and elevator microswitches RMS,AMS and EMS all closed, a circuit is established to engage switchholding solenoid 12. Thus, should any malfunction occur which wouldcause any of the servo clutches to separate, the affected microswitchvwill open, deenergizing engage switch holding solenoid 12 and allowingthe engage switch 9 to drop to either its yaw damper position or offposition, depending upon which servo experienced the failure. The engageswitch is so designed that, should the malfunction not affect the yawdamper elements, the engage switch will automatically drop to yaw damperposition and remain there. This is accomplished in part by separateenergization of the engage switch holding solenoid 12 through yaw dampercontact on switch 1t).

Closure of relay K-IGZA also serves to place roll command computer 208and pitch command computer 308 in their engage positions wherein theywill respond to any input commands thereto. Through the operation of K-102A, roll integrator 11 is also rendered responsive to any input signalapplied to its input by closure of switch K- 192A-3. The tachometergradients of the roll command computer 268 and the pitch commandcomputer 368 are adjusted for their normal autopilot operation throughener- 9 gization of relays K-lt (FlG. 12) and K-190 (FIG. 7) and closureof engage relay switches K-ltZB-Al and K- NZD-1, respectively.

It has been described above that the servos cannot be engaged if theaircraft is in a banked attitude of greater than plus or minus 35. Ifthe autopilot is engaged at bank angles less than the craft willauomatically roll to level flight and the steady state heading, i.e.,the compass heading, at engagement will be maintained. This isaccomplished as follows. Since during the disengage mode or yaw dampermode the roll computer 29S followed up on the vertical gyro signal, theroll attitude of the aircraft exists in synchro 2l9 (FlG. 1B) and thesine of this bank angle is present in resolver 222, also positioned bythe roll command computer shaft. Therefore, upon engagement throughswitch K-ltZA-, the stored bank signal is applied to the input of theroll command computer and drives the latter to zero the signal, and inso doing causes aileron operation through gyro synchro 213 and rollcommand computer synchro 219 to roll the aircraft to a level attitude.

Referring now to FIGS. 5 and 1B, a pair of liquid level switches 223 areprovided and are responsive to lateral accelerations of the craft suchas would be produced by a bank angle of about 5. These switches arearranged on the gimbal of the roll gyro as represented by the dottedline connection in FIG. 1B and are normally maintained level by gyrorigidity and hence are very sensitive to lateral accelerations. Theseswitches serve to close a circuit from the instrument circuit breakerthrough a suitable transformer and bridge rectier 224 to ground, thusenergizing erection cut-off relay K-ENM, energization of which serves tocut-off the erection on the vertical gyro until the bank angle of thecraft is less than 5 roll, i.e., the lateral acceleration sensed by theswitches drops below a predeterined value. Energization of erectioncut-off relay K- 3M- also energizes bank threshold relay K-lli throughaction of switch K3d41 (FIG. 6). If, however, the bank angle is under 5gyro erection stays on and the bank threshold relay l-1tl remainsdeenergized. Referring now to FIG. 1l, it can be seen that even with thepilot engage relay closed, yaw comand computer 16S is clamped throughthe deenergization of bank threshold relay switch K-lt-, therebyallowing yaw command computer relay K-17t) to be energized, whichdisconnects the output of the yaw followaup amplifier with its follow-upmotor. Also, as shown in FIG. 1l, gust clamp relay K-lll is alsoenergized, thereby supplying a heading reference signal to the rudderservo system through gust clamp switch K-lll-l and to the roll commandcomputer through gust clamp switch K-fill-Z. In the foregoing manner, ifthe pilot is engaged at less than 5 bank angle, the craft will rolllevel and maintain the heading obtaining at engagement.

lf the autopilot should be engaged at greater than 5 bank angle, thecraft will make a coordinated maneuver to level flight and maintain theheading achieved when the bank angle drops below 5. This roll to levelflight is accomplished in the same manner as before, but while theaircraft is at a bank angle greater than 5 the bank threshold detectorK-Htl will be energized due to the cnergization of erection cut-offrelay Kitii as a result of closure of one or the other of liquid levels223. Energization of the bank threshold detector opens the switchK-llQ-l and deenergizes yaw computer clamp relay K- 179, placing yawfollow-up 168 in follow-up on the signal from the directional gyro 167.Simultaneously, gust clamp relay K-i is likewise deenergized so that noheading reference signal is applied to the rudder servo nor to the rollcommand computer Ztl. Thus, as the roll command computer 268 reduces thebank angle signal in resolver 222 toward zero, the directional gyroremains in follow-up until the bank angle drops below 5 at which timethe yaw command computer 108 is clamped as described above.

Upon engaging the Hight control system, the pitch attitude existing atthe time of engagement is maintained and the craft may be returned tolevel flight by operating pitch command knob 16 or by selecting thealtitude control mode. If the autopilot is engaged at 0 pitch attitudewith the altitude control olf and the pitch knob in detent, pitchfollow-up clamp relay K-191 (FIG. 7) is energized and the pitch commandcomputer 308 is clamped. The output of the pitch command computersynchro 317 is zero at this time by virtue of its following up on thevertical gyro signal during the disengage mode and any deviations fromlevel night after engagement will cause the craft to be brought back tolevel flight due to the direct connection from vertical gyro pitchsynchro 318 to the elevator servo system 304. Also, if the autopilot isengaged in, say, a pitch down attitude, no pitch signal will be suppliedto the elevator servo system, again due to the follow-up action of pitchcommand computer during the disengage mode.

The autopilot is now fully engaged Iand is conditioned for acceptance ofany maneuver command signals generated either by manually insertedcommands through the turn and pitch knobs 14, 16 or through externallyinserted commands, i.e., radio commands or altitude control commands.

Autopilot Short Term Stabilization With the automatic pilot engaged,short term attitude stabilization is provided by the yaw, roll and pitchangular -accelerometer signals produced by the yaw, roll and pitchaccelerometer pairs. The function of the roll and pitch accelerometerpairs in stabilizing the craft against short period disturbances isexactly the same as l@hat previously described with respect to the yawchannel operating the yaw damping mode and a detailed discussion thereofwill not be repeated. As stated, the upper and lower roll accelerometersignals are combined so as to produce a resultant signal proportional tothe angular acceleration of the craft, the signals being modified aspreviously described `and applied to the aileron servo system 2534 wherea surface rate proportional to the measured acceleration is produced tothereby stabilize the craft against short period roll disturbances. Thepitch angular accelerations are suppressed in the same manner. However,as in the yaw channel, the accelerometer 106 mounted at the C.G. of thecraft supplies a separate signal proportional to the verticalacceleration thereat. It will be noted in FIG. 1C that the details ofthe lead filter and lag-lead filter are shown in det-ail, `and it shouldbe understood that the corresponding lters in the other accelerometerchannels are substantially identical, the functions thereof having beenpreviously described in connection with the yaw damper mode ofoperation.

Manual Maneuver Commands Assume now that the craft is flying straightand level and it is desired to manually maneuver the craft about itspitch axis. For pitch maneuver commands by the human pilot, pitchfunction selector knob 18 must be in its Pitch Knob position (FIG. 4).With the pitch function selector knob 18 in this position, the pitchattitude of the craft may be changed by rotation of spring centralizedpitch rate command knob 16. Rotation of this knob out of detent closesswitch 17 and energizes relay K-lto (FIG. 3) which, in turn, deenergizespitch follow-up clamp relay lli-191 (FIG. 7) by opening of switch K-lo-Sto thereby unclamp pitch command computer 363 (FIG. 13) by closing ofswitch K-191-1. At the same time, a pitch command signal proportional tothe sense and magnitude of the displacement thereof is generated throughpotentiometer 314, the magnitude and sense of this signal being in turnproportional to the desired sense and rate of change of pitch attitudeof the craft. This signal is applied through now closed pitch detentrelay switch K-l-Z to the input of pitch followup amplifier 3tl8-whichenergizes pitch follow-up motor 315 so that it drives in the directionand at a rate proportional to the direction and magnitude `of the pitchcommand signal by virtue of the speed feedback voltage from generator316. It will be noted that in this mode relay K-190' (FIG. 7) isenergized and relay switch K-190-1 is moved to its high tachometergradient position, thereby decreasing the rate at which the follow-upmotor can operate as compared with its rate in the disengage mode, thatis, to a rate consistent with craft response characteristics to inputcommands. It will be further noted that the motor 315 will continue todrive until the pitch command signal is reduced to zero as by manuallyreturning knob 16 to detent. Rotation of the pitch command cornputersynchro 317 with respect to vertical gyro synchro 318 will produce anerror signal which is supplied to the elevator servo system 394 tothereby produce a pitch rate of the aircraft proportional to such error.In order that the elevator deiiection producing the commanded rate isnot initially yopposed by the angular pitch acceleration which normallytends to oppose any rotation of the craft about its pitch axis, the rateof change of the pitch comm-and signal as measured by follow-upgenerator 316 and hence a pitch acceleration term is applied throughswitch K-11`23 (deenergized in the manual pitch command mode) and asuitable isolation transformer 319 to the input of amplier demodulator320 in the angular acceleration output channel in opposition to theangular acceleration signal, thereby bucking o-ut any pitch angularacceleration signal which would oppose the pitch command.

When the desired pitch attitude of the craft has been lachieved, thepilot releases or re-centers the pitch knob 16, thereby zeroing thepitch command signal from poten tiometer 3,14 and closing pitch detentswitch 17. Thus, pitch detent relay K-106 is deenergized and Isl-191energized yto thereby clamp pitch command computer 308 at the positionit then had. The vertical gyro thereafter stabilizes the aircraft at thenewly established pitch attitude. The craft may be returned to levelflight attitude `by an opposite sequence of operation of the pitchcommand knob 16, or, if desired, by rotating pitch function selectorswitch 18 to the ALT position. The altitude control mode will bedescribed below.

As stated, two types of manual turns may be Iaccomplished with theautopilot of the present invention; by manual turn rate commandsinserted through operation of turn rate controller 209 and bypreselected heading commands through heading selector 210. The selectionof these modes is accomplished through the pilots turn function selectorswitch 19 to either the Turn Knob or HDG SEL positions, in each of whichinterlock circuits are established whereby command turns from thesesources are supplied to the autopilot.

As shown in FIG. 4, the turn knob position of function selector switch19 is the normal position, the knob being spring centered to thisposition. With the knob 19 in this position, relay K-140` lis energized,conditioning the system for Iadditional operation in either of the radiomodes to be hereinafter described. Turn rate commands are institutedthrough turn knob 14, rotation of which out of detent position closesdetent switch 15 and energizes relays K-104A and K-104B. Simultaneously,a voltage is generated across potentiometer 225 which is proportional tothe magnitude and sense of such displacement, which voltage is in turnproportional to the magnitude and sense craft rate of turn which it isdesired to make. Energization of relay K-104A serves to unclamp the yawcommandV computer 108 through the opening of switch K-104A-1 (FIG. 11)with the resulting deenergization of yaw computer clamp relay K-170 andgust clamp relay K-111. Thus, as the aircraft turns in response to theturn command signal, the yaw command computer 108 follows up ion .anysignal from the directional gyro 107. Energization of relay K-104Bcloses switch K-104B-4 (FIG. which serves immediately to cut olf theerection controls of the vertical gyro and any slaving controls of thegyromagnetic compass system of which the directional gyro 107 randheading selector 210 may form a part. Erection cut-oli relay K-304 also,through switch K-3041, energizes bank threshold relay K-110 (FIG. 6)which, in turn, opens switch K-110-1 in parallel with the gust clamprelay switch K-111-1 in the yaw command computer clamp and gust clampcircuit (FIG. ll). Function of'these switches will be describedhereinbelow. Relay K-104B also closes switch K-104B-2 (FIG. 1B) whichconnects the turn command signal to the roll cornmand computer 208.

The turn rate command signal applied to roll command computer 208produces a rotation of roll computer motor 226 at a speed determined bygenerator 227 and through an angle proportional to the sine of the bankangle required for the rate of turn commanded. Synchro 219 thus biasesor shifts the bank angle reference provided by synchro 218 on thevertical gyro 6 through an angle in accordance with the rate of turncommand. The signal from synchro 219 is applied to the aileronservomotor and causes the craft to bank to that bank angle and thereforeto turn at a rate corresponding to that commanded. Of course, shortperiod roll stabilization continues to be supplied by the upper andlower accelcrometers 295 and 206. Any initial opposition to the rollcommand by the roll accelerometers and the inherent lag in servoresponse produced thereby is not objectionable in this channel. As willbecome evident later on, this lag is actually accentuated in other modesof operation. As explained above, the roll acceleration signal isapplied through yaw damper integrator 124 to apply rudder in a sense tooppose any adverse yaw due to the roll rate. As the craft turns inresponse to the bank angle, the yaw command computer 108 follows up onthe directional gyro 107. However, since gust clamp relay has closedswitch K-111-1, any heading error signal exceeding the follow-upcapacity of the yaw command computer 108 is applied to the rudder servosystem to thereby provide heading stabilization during the turn. Shortand long term turn coordination is supplied through the C.G. mountedaccelerometer 106 operating directly into the rudder servo system andthrough electronic integrator 124 into the rudder servo system,respectively.

In order to prevent any loss in altitude during the turn, due to thedecrease in vertical lift of the wings by banking of the craft, a liftcompensation signal dependent upon bank angle is inserted into the pitchor elevator channel of the autopilot. It can be shown that the precisemathematical value of this compensating signal may be very closelyapproximated by the expression (1-cos g5) where is the bank angle. Thus,a further winding on resolver 222 rotated by the roll command computerm0- tor 226 provides a signal proportional to cos qb. This signal isapplied as an input to a (l-cos qb) computer 323 where it is combinedwith a fixed constant voltage. The resultant signal is applied throughtwo branches to the pitch channel of the autopilot through a suitableisolation transformer 324. In one branch the signal is parametercontrolled as at 326 in accordance with the reciprocal of dynamicpressure q since the value required for lift compensation varies withair speed. This lift compensation signal is applied to the input of theelevator servo system 304 and produces an elevator deflection such as totend to maintain the altitude of the craft constant during the turn.

When it is desired to stop the turn, the pilot rotates the turn knobback to its detent position, the craft mmediately rolling to leveliiight and, through interlock circuits, it will ily to and maintain theheading obtaining when the bank angle has decreased to below 5 Return ofturn knob 14 to its zero position opens detent switch 15 and deenergizesrelay K-104A and K-104B; K-104B-2 and -B-4 serving, respectively tosever the connection between the turn command Ztl? and the input to theroll command computer 2tlg (FIG. 1B) and to place the erection controlrelay K-Stlft under the control of liquid levels 223 (FIG. 5). Also,K-lll-/A-l establishes a connection from the D.C. supply to bankthreshold relay switch K-lltl-l and gust clamp relay switch K-lll-l(FIG. ll). Therefore, as soon as the bank angle drops to 5, erectionrelay K-Ztld will become deenergized due to the leveling of the liquidlevels 223. At the same time, bank threshold relay K-llfl will be alsodeenergized, again deenergizing yaw command computer clamp K-17f andgust clamp K-lll. Thus, the yaw command computer 163 is again clampedand the directional gyro 107 provides it heading stabilization signal tothe yaw and roll channels of the autopilot through gust clamp relayswitch K-lll-Z (FIG. 1B) to maintain the craft in straight and levelflight at the heading then obtaining.

If when the craft is rolling out of the bank maneuver and the bank angledrops below 5, but a sudden gust should cause the craft momentarily toincrease its bank angle over 5, the switching sequence above describedwould again unclamp the yaw command computer causing to follow up on anydirectional gyro signal. Since a gust is normally of generally shortduration, such momentary operation of yaw command computer isundesirable and means have been provided for allowing the yaw commandcomputer to be unclamped only upon commanded turns rather than upon agust produced bank angle.

'Upon gust clamp relay K-lll becoming energized, switch K-lll-l (FIG.1l) in parallel with bank threshold switch K-11B closes, maintaining adirect D.C. path to the `gust clamp yand yaw command computer clamprelays, even though bank threshold relay K-llf) is energized due to agust detected by liquid levels 22.3.

The craft is now back in straight and level flight, holding the headingdefined by the directional gyro 1%7 and an altitude defined by verticalgyro 6. It will be noted in IFIG. 4 that the manual turn knob mode takesprecedence over any other mode of operation of the autopilot through theuse of a suitable electromechanical holding latch 227 coupled withswitch 19, the energization of which is controlled in dependence uponwhether or not turn knob 14 is in or out of detent. For example, if knob19 were in any position other than Turn Knob and the pilot turned turnknob 14, K-ltldB would become energized, closing switch K-tMB-S in FIG.4 and energizing holding switch solenoid K-Zl, automatically centeringselector knob 219 through its centralizng spring.

Command turns may likewise be made through the pilots magnetic headingselector 216. The details of such a selector are shown in theabove-identified parent application and particularly in FG. 5 of thatapplication. Through this type of turn control, the pilot may turn toany selected magnetic heading and the craft will smoothly bank up into acoordinated turn and then roll to level flight at the magnetic headingso selected. The normal procedure for making a heading type turn is toselect the heading it is desired to fly on the heading selector 210 withthe turn function selector knob t9 in its normal turn knob position. Thesignal so generated will wind up on open contacts of switches K-l1-lland 2 (FIG. 1B). Then, when it is desired to make the turn, the pilotmerely turns the turn function selector knob to the HDG SEL position,the latter setting serving to institute the turn of the craft. In thismanner, the desired heading may be preset and the actual turn of thecraft instituted at any desired time thereafter, a feature which may behighly desirable, especially when holding in a stack preparatory tolanding. Alternatively, the turn function selector knob may be set toheading select and the heading then selected on the selector 210, thecraft smoothly following the heading as it is being selected.

As described in the parent application, setting of the selector 210generates a signal in a synchro coupled with the heading indicator whichis proportional to the difference between the actual heading of thecraft and the desired heading so set. Switching of turn selector knob 19to the heading select position energizes K-ltll (FIG. 4) and suppliesthe heading error signal of heading selector 2l@ to an amplifier limiter228 through switch K-ltll-l (EKG. 1B). Also, switch K-lfll-2 closes acircuit between heading selector Zltl and the roll integrator l1 Whoseoutput is combined with the heading error signal and applied to theinput of amplifier limiter 228. The integrator serves, in the headingselect mode, to prevent the aircraft from turning due to any persistenttrim error which is reflected in a persistent heading error. It will benoted that the heading error signal from heading selector 2l@ is firstpassed through a transient integrator 229 for smoothing out any abruptmotion of the heading selector knob. Energization of relay K-ltll alsoarms a smooth engage circuit 23d` through switch K-ltlll-B, relay K-lil(FIG. 13) and its associated switch K-llS-ll (FIG. 1B). This smoothengage circuit functions through a capacitorcharge time type circuit -tolimit the speed with which the heading error signal is allowed to beapplied to the roll command computer 2tlg. This is especially desirablein the above-described preferred manner of making a heading select typeturn when there might otherwise be a switching transient.

Inasmuch as the magnitude of the error Signal from heading selector 2idmay be Very large; as for example when, say, a 1801" turn is selected,and because a bank angle proportional to heading error is provided tothe aileron servo system, the heading command signal is limited inlimiter 228 to a magnitude such as not to command an excessive bankangle. As illustrated in FIG. 1B, the limits imposed on the bank anglemay be varied, depending upon the mode of operation of the aircraft.These will be discussed later in connection with the radio guidancemodes. The remainder of the operation of the autopilot in the headingselect mode is exactly the same as that in the turn knob mode,particularly in regards to the yaw command computer clamping andunclamping and gust clamp operation.

Flight Path Control In this section of the present specification will bediscussed the operation of the autopilot under the influence of commandsgenerated in accordance with departure of the craft from preselected orpredetermined flight paths, that is, flight paths that are defined bysuch references as barometric data and navigational or other radiobeams.

If it is desired that the autopilot control the aircraft to maintain adesired pressure altitude, function selector switch 18 is turned to theALT position. As shown in FIG. 4, such switching will energize altitudecontrol relay K-dl which serves to energize altitude control 310. lnpractice, relay K-lll is in the altitude control unit 31d and functionseffectively to clamp a barometric altitude reference member, such as ananeroid bellows, in the position it had when clamped and any error inthe altitude thereafter will produce a signal output from a suitablepick-off device proportional to such deviation. Switch 18 also energizesa second altitude control relay K-lll which, in turn, closes switchlli-M54 (FlG. 1B) thereby supplying the altitude error signal to theautopilot command circuits. Relay K-llS also controls switch K-ll- 2(FlG. 9) which serves to energize vertical path relay K-ll2- Operationof this relay in turn opens opens switch K-llZ-l (FG. 7) to therebydeenergize pitch follow-up clamp K-ltl and allow pitch command computer39S to be responsive to any signal applied to its input amplifier.

At this point it should be stated that if the craft is in altitudecontrol and the pilot wishes manually to change attitude, he need merelyrotate the pitch command knob 16. This action will again close pitchdetent relay K-ltlo which, in addition to putting in the pitch commandas de- Vof elevator servo system 3M.

scribed above, also deenergizes a mechanical interlock relay K-ZG()which automatically returns the function Selector switch knob 18 to itspitch knob position. Conversely, if the aircraft is in a climb or divethrough pitch knob operation and it is desired to level out at aparticular altitude, the pilot need merely switch the function selectorswitch 1S to its ALT position, such switching making the connectionsdescribed above to hold the altitude at which the switch 13 wasoperated. The details of the altitude control system are disclosed inthe above U.S. application Serial No. 571,788.

As will become evident, when the craft is placed in the altitude controlmode, the signals controlling the elevator servo system include analtitude displacement error signal, a damping term derived from theintegral of vertical acceleration, and an integral control .term derivedthrough integration of any persistent altitude error. In order for suchintegral terms to be derived, certain modifications in the pitch commandcomputer 368 must be accomplished. It will be remembered that in thepitch maneuver command mode, the pitch command computer followed uprapidly on the pitch rate command signal. In the altitude control mode(as well as the glide slope mode to be described later), pitch commandcomputer 398 has its gain changed to such an extent that it may operateas a long term integrator. Thus, when the altitude mode is selectedthrough selector switch i8, certain interlocks are effected to makethese changes in the pitch command computer. With selector switch 13 inits altitude position, altitude relay K-llS is energized through engagerelay K-lZC-Z (FIG. 4), thereby energizing vertical path relay K112through engage relay K-1tl2D-1 (FIG. 9). Energization of K-112 opensswitch K-1121 (FIG. 7), thereby deenergizing relay K-191 and unclampingthe pitch command computer 308. Since the pilot is engaged, K-19tl isenergized thereby increasing the amount of rate generator feedbacksignal to the input of the amplifier of the pitch command computer.Energization of K-112 also closes switch K-llZ-Z (FIG. 8) which, inturn, energizes solenoid MiG-101. This solenoid functions to shift gearsin the gear transmission 322 between pitch command computer motor 315and its output synchro 317. In the manual pitch rate maneuver mode, ahigh speed gear Yreduction (for example 5000A) is required so that thepitch command computer follows up rapidly on the pitch rate commandsignal. However, to perform as an integrator, the gearing between thepitch command computer motor 315 and output 317 must be a low speedgearing such that considerable rotation of the motor 31S will produceonly a small rotation of output synchro 33.7

(for example l50\0{):l). Thus, operation of the selector knob 118 to itsALT position not only switches in the altitude control 310 but also,through the interlocks just described, conditions the pitch commandcomputer 368 to operate as a long term integrator of any persistentaltitude error.

For ythe purposes of the present application, the altitude cont-rol 310supplies an output proportional to the displacement of the craft fromsome reference barometric altitude, this signal being applied directlythrough an amplifier, lag and lim-iter circuit B2i directly to the inputThe limiter serves to limit Ithe magnitude of the pitch attitudecommandable by the altitude error signal. This displacement signal 1sbucked by a signal from the vertical gyro synchro 31S 'through synchro317 on the pitch command computer Sii/8 in a suitable summing circuit.This signal provides short term altitude control which corrects forgusts and other disturbances. The lag serves to smooth the altitudeerror signal. Inertial path damping is provided in the present autopilotby a signal proportional to normal accelerations of the craft, i.e.,parallel to direction of gravity, as detected by the CG. or verticalaccelerometer 106 (FIG. lC), this signal being appiied through ahighpass or lead filter,` to the input of the pitch command computer3tlg where it is integrated and applied to the pitch servo system as analtitude rate term. However, with the airplane in a turn, the CG. orvertical aocelerometer 106 would sense `the centrifugal accelerationproduced by the :turn `and therefore would tend to nose the craft downin order lto reduce such acceleration. Since the vertical accelerometeerobviously cannot sense the difference between normal acceleration and aIvertical acceleration compo-nent due to centrifugal force, the otherbranch of the output of the l-cos qb computer is applied in the outputof vertical accelerometer lila in such a sense as to cancel only thatcomponent of vertical acceleration produced by turning, i.e., thecentrifugal force component. However, since the li-cos qb signal is anapproximation, a small vertical acceleration signal may exist at highbank angles. The difference between the normal acceleration signal andthe l-cos 15 signal is therefore applied to the integrator or pitchcommand computer 308 through a highpass 0r lead filter to thereby.remove any longer period or steady state acceleration signals due tothe difference between the actual and computed normal acceleration. 4I-twill be noted that the normal acceleration or inertial path dampingsignal is automatically inserted into the autopilot whenever a verticalpath mode is selected, i..e., when altitude is switched on or when aglide slope mode is Irendered operative. compiished through switchisi-11.243 (FIG. 1C), closed by energization of vertical path relayK-IIZ (=FIG. 9).

Furthermore, the displacement signal from the altitude control 31) isalso applied to the input of the pitch command computer 33,8 which, inthis mode, serves to integrate the same to thereby allow any persistentaltitude error to go to zero. In other words, integration of thealtitude displacement signal through pitch command computer 30S allowsthe altitude control mode to be disengaged without a transient.

`lt will be understood lthat the operation of the pitch channel of theautopilot in response to an error signal rom a radio beam, such as anILS glide slope beam, will be exactly the same as in the altitudecontrol mode, ie., that the elevator servo system is controlled inaccordance with a glide slope displacement signal, the integral ofnormal acceleration -or the rate of change in altitude which, in theglide slope, is the same as the rate of change of glide slope error, andthe integral of any long term glide slope error signal.

The automatic pilot of the present invention may be controlled toautomatically seek, approach, and thereafter maintain a iiight pathldefined by radio signals. Radio beam guidance facilities fall generallyinto two categories, VOR facilities and ILS facilities. As is known, theformer provide on route navigation beams while the ILS, of course,provide terminal area or instrument landing radio guidance beams, thelatter including overlapping beams for providing localizer guidance inthe horizontal plan-e and glide slope guidance in the vertical plane.

As in the other modes of operation of the autopilot, selection of thedesired mode of operation controls interlock circuits whereby the pilotis conditioned for operation in such mode. The following discussion ofthe radio guidance mode will be divided into three sections; VORcoupling, localizer coupling, and finally glide slope coupling. It willbe noted that in FIGS. 4 and 14 certain of the switching and relaysschematically illustrated in FIG. 4 have been repeated in FIG. 14, andin the following description reference may be made to either figure forthose elements common to each.

When the automatic pilot is rst engaged, the turn selector function knob1'9 is automatically in its Turn Knob position as above described. Undersuch conditions, D.C. power is applied to lateral beam sensor relay K-A,closing switches K--1 and K-140-2, the latter preventing energization ofradio beam coupler on relay K-103. With the selector knob 19 in the TurnKnob Such switching is acl? tion upon subsequent turning of the selectorknob to the VOIR/LOC position.

Assume now that it is desired to approach and maintain a VOR beam. Priorto operation of turn selector knob 19, the pilot selects the desiredomnirange frequency as by frequency selector `493 and sets in, throughomni-heading selector 464, the bearing of the omniradial he desires toily. Such `omni-heading selector is decribed in detail in U.`S. Patent2,732,550, which is assigned to the same assignee as the presentapplication. -As is shown in that patent, the omni-heading selector b4provides an output signal which is proportional to the angular deviationof the aircraft fro-m the omniheading. Tha-t is, the signal isproportional to the angle between the instantaneous heading of theaircraft and the bearing of the omniradial. Of course, through hisnormal navigation facilities, he may maneuver the craft through the turnknob 14 to a position, as determined by navigation charts, etc., toi thevicinity of the selected omnistation. When he desires automatically toapproach and thereafter maintain the selected radial he rotates turnfunction selector knob l@ to the VOR/LOC position. As shown in PlC-S. 4and l4, the selector' switch i9 is a make before brea type switch sothat such switching will not deenergize the lateral beam sensor relayK445i.

Before describing further the interlock circuitry of FIG. 14, portionsof the beam coupler 400 (FIG. 1A) should be described. Lateral radioreceiver 2li provides a signal proportional to the lateral displacementof the craft from the selected radio course. This signal is modulatedand amplified at 4ii5 and applied as an input to a position follow-upservo loop liti-6 which, through feedback connection from potentiometer4d?, positions the follow-up shaft 46S, through amplifier 409, motor419, and rate generator 411, to a position corresponding to such lateraldisplacement. On the output shaft 49S of the approach coupler follow-upservo 4de is the wiper of a sector switch 412. This sector switchcomprises two conducting segments, one of which is relatively long andthe other fairly short. The approach coupler follow-up loop 4% is sodesigned that when the aircraft is located at a distance from the beamgreater than a first predetermined distance (for example, the distanceat which the beam displacement signal is greater than 155 milliamps.),the wiper of the switch 412 will lie on the large sector appropriatelylabeled plus or minus 155 pa. Similarly, the small sector of the switchrepresents a second predetermined lateral distance from the beam (forexample, a distance represented by a displacement signal ofapproximately plus or minus 50 milliamps.), so that when the beam iswithin said predetermined distance from the beam center the wiper willlie on the short sector. This short sector is appropriately labeled plusor minus 50 ha. The nonconductive portion between the ends of the twosectors therefore represents the distance or displacement of the craftfrom the beam represented by a radio signal having a magnitude betweent) ha. and 155 tra.

At this time it will be pointed out that the gain of the radio signalfollow-up loop 466 may be changed under varying beam coupling conditionsas by changing the magnitude of the rate feedback signal from generator411. It should also be mentioned suitable stops on the shaft 49S areprovided so that the motor cannot drive the sector switch wiper throughan angle greater than 360 in any one direction. Such stops may, forexample, be included within the potentiometer 497, suitable clutch meansbeing provided for preventing damage to the servomotor. Now assume thatit is desired automatically to approach and maintain the radio beamselected and also that the craft is at a distance greater than thatrepresented by a 155 pa. displacement signal. Also assume that the crafthas been placed on a heading which will cause it to intercept the beam.Switching of function selector switch knob i9 to VOR/LOC position merelymaintains lateral beam sensor relay Kl4t) energized through holdingswitch K-lltl-l with the sector switch wiper, of course, being on thelarger sector, as illustrated in FIG. 14. As the craft approaches theedge of the beam, the sector switch arm begins to move towards the endsof the contact sector switch 412. As the craft approaches the beam andthe displacement signal drops to tra., the wiper of switch 412 leavesthe 155 lua. contact sector, thereby deenergizing lateral beam sensorrelay K-lnl@ which, in turn, allows switch K-l4tl-2 to move to itsdeenergize position, thereby energizing radio coupler on relay K-lltit.With the energization of relay K-lti, switch K-ltlS-l (FIG. 1B) closesand applies the sums of the signal proportional to displacement of thecraft from the beam and a signal proportional to the heading of theaircraft with respect to the beam to roll control channel of theautopilot. The beam displacement signal is taken directly from theoutput of the modulator pre-amplifier 405 through an amplification stage413 while the heading signal is obtained from the omni-heading selector464. The algebraic sum of these signals is applied to the input ofamplifier limiter 22S in the input to the roll command computer Zilwhere the combined signals command a bank angle proportional thereto,just as in the case of a turn rate command signal.

ln the limiter circuit, the radio and heading signals are limited so asto limit the bank angle commanded thereby. With the energization ofradio coupler on relay K-l, the coupler may be said to be in a bracketmode, and in this mode the limits imposed by limiter 223 are such thatrelatively large bank angles may be commanded thereby enablingbracketing to occur swiftly. Also, the smooth engage circuit 23d isrendered effective through smooth engage relay lli-119 (FIG. 13) andswitch K-llS-l controlled thereby as in the heading select mode, thisbeing accomplished by means of switch K-liBS-S. Smooth engage circuit23th is essentially a condenser network which serves to allow the rollcommand signal from the radio coupler to slowly build up over apredetermined time period from zero to the maximum command value. A timeconstant of about four seconds has been found to be satisfactory forthis buildup. The operation of the automatic pilot in response to thebank angle command is the same as its response to a command either bythe turn rate knob 14 or the heading selector 2li) and will not berepeated.

Since the craft is being controlled in accordance with the displacementfrom the beam and the rate of approach as determined by the heading ofthe craft toward the beam, an asymptotic approach path will be executed.As the approach continues, the arm of sector switch 4t2 will approachthe edge of the short contact ector and when the craft is at a distancerepresented by a signal strength of approximately 55 milliamps., thecontact arm will contact this sector. When this occurs, on course relayK-M will be energized provided that relay switch K-l44-l is deenergized.The latter switch is controlled from cross-course velocity relay K-144which is energized when the cross-course velocity of the aircraft asdetermined, in the present VOR coupling mode by the magnitude of thesignal from the omiheading selector which, of course, is a measure ofthe rate of approach of the craft with respect to the beam. lf, for somereason, the cross-course velocity exceeds a predetermined value, forexample, a velocity resulting from a difference between the heading ofthe craft and the bearing of the beam of, say, 15, determined bycross-course velocity sensor 414 which is any suitable signal magnitudesensing circuit, such as an amplifier suitably biased, the relay K-l-/itwill not deenergize and the on course relay will likewise not beenergized, thereby leaving the craft in the bracket condition. Thus, ifthe initial heading, air speed, etc., conditions could not result in anasymptotic approach, the on course 1.19 mode will not be instituted andthe craft may be allowed to go through an overshoot and continue thebracket until, on the next approach, cross-course velocity falls belowthe threshold value as determined by sensor With K-144 deenergized, oncourse relay 141-145 becomes energized. Energization of the latter relaywill, through switch K-143-3 (FIG. 1E) bypass an attenuator to therebyincrease the effective gain of the beam displacement signal and, throughswitch K-143-2, add the integral of the beam error signal through rollintegrator 11 whereby to more tightly hold the craft on the beam and tocompensate for the effects of steady crosswinds tending to blow thecraft off course, respectively. Also, on course relay K-143 changes thelimits imposed on the bank angle command as by limiter impedanceadjustment through switch K-143-6. This change is in a direction todecrease the bank angle commanded by the radio and heading signals sincesmaller heading changes are required to maintain the beam once the crafthas acquired the beam. As a further advantage of such bank limiting, asmoother and more comfortable ride is achieved.

The craft will be maintained on the selected VOR radio through jointeffects of the radio displacement, heading and bank angle signals. Asthe craft approaches the VOR transmitter, the displacement signal willbecome erratic in a region directly over the transmitter, this regionbeing known as the zone of confusion because no suiiiciently Welldefined radio track information is available for control purposes inthis region. Therefore it has been found desirable to sever completelythe control of the aircraft through the radio signal and to leave onlythe heading signal operative to control the craft. This over-the-stationcontrol is disclosed in detail in U.S. Patent No. 2,881,992, which isassigned to the same assignee as the present invention. However, a briefdescription will be included herein for the sake of continuity ofdisclosure.

In FIG. 1A, it will be noted that when on course relay K-143 isenergized (FIG. 14), relay switches K-143-4 and -5 are actuated, theformer serving to connect the output of rate generator 411 to the inputof an over-thestation sensor ampliiier 414 and the latter serving toconnect the output of this amplifier to the input of radio beamamplifier 413. As the displacement term becomes erratic, its rate ofcomponent as sensed by rate generator 411 will become even more erratic.The overthe-station sensor 414 comprises a circuit which is responsiveto a predetermined magnitude of signal applied to its input and willoperate to supply an output signal of a predetermined magnitude when theinput voltage exceeds its predetermined magnitude. The output of theover-the-station sensor 414 is employed to effectively cut off theoperation of displacement signal amplifier 413 such as by biasing theamplifier to cut-olf or by other means to thereby effectively remove theradio displacement term from the control system during this period oferratic radio signals. The craft will therefore be caused to maintainthe heading it was on at the time the overthe-station sensor becameeffective. As the craft comes out of the zone of confusion, the ratecomponent of the displacement signal will drop to a low value, and aftera predetermined time period thereafter, say, 4 seconds, the displacementamplifier 413 will once again be rendered effective to supply the radiodisplacement term to the autopilot for continued control of the craft asit continues outbound on the reciprocal VOR radial under the control ofthe radio displacement, heading and bank angle signals.

If it is desired to use an ILS approach facility, the turn functionselector knob 19 should be returned to the Turn Knob position, thuspreparing for pilot-inserted maneuvers usually necessary to arrive inthe vicinity of a desired ILS localizer beam. As will become apparent,such operation will also avoid any transients which might otherwiseoccur when tuning to an ILS frequency. When the pilot tunes hisfrequency selector 403 to an ILS frequency, as indicated schematicallyat 415, a pair of ILS relays iti-126 and .ii-145 are energized, therebyconditioning the autopilot for an ILS approach. The approach to alocalizer beam is very similar to that to a VOR beam except that in thelocalizer mode the heading is not used as the beam damping or rate ofapproach term, but the actual rate of approach of the craft asdetermined by the rate of change of the radio displacement signal isused. This signal is derived at the output of rate generator 411 in theradio displacement follow-up loop 496.

Energization of ILS relays K-126 and K-145 through the selection of anILS frequency serves to energize switches K-126-1 and K-126-2 whichrespectively supply a beam displacement signal, suitably attenuated forlocalizer bracketing and a radio rate signal, also suitably attenuatedfor a localizer bracket, which two signals are combined and supplied tolateral beam amplifierV 413, the output of which is supplied to theautomatic pilot turn command channel. It will be noted also that ILSrelay K-llZi-S serves to remove the heading signal when in the localizermode.

As in the VOR coupling mode, if the craft is considerably displaced fromthe localizer beam, i.e., in excess of 155 tra., lateral beam sensorrelay ISI-140 will be energized and switch K-14il2 will keep radiocoupler on relay K-163 deenergized. Assuming that the craft is on aheading which will cause it to intercept the beam, when the beamdisplacement signal drops to a value below 155 ita., beam sensor relay{(-140 will become deenergized, thereby energizing radio coupler onrelay K-lii as before. As seen in FIG. 1B, the radio coupler on relayswitch K-103-1 supplies the radio displacement plus rate signal toamplifier limiter 2.28 to the input of the roll command computer therebyto cause the craft to roll to an angle determined by the limited radioplus radio rate signal. The resultant radio displacement, radio rate,and bank angle signals command an asymptotic approach of the craft tothe beam depending upon how far out the approach was initiated. Thecraft may and very likely will reach a displacement from the beamrepresented by a 50 fia. radio signal and the beam sensor switch armwill contact the 5G pa. sector. However, the craft will continue theapproach with no change, i.e., the final approach parameters will not beengaged, until the approach relay K-142 (FIG. 14) becomes energized.This can occur only under certain conditions. ILS frequency selectorrelay K- fulfills a first condition by operation of relay switchK-145-1. The second condition is that the cross-course velocity must bebelow a predetermined value, say 2 millivolts per second, therebydeenergizing cross-course velocity detector K-144. The third conditionis that the turn knob function selector must be in its Glide Pathposition and the glide slope beam intercepted.

ILS frequency relay K-145 also energizes switch K-14S4 (FIG. lA) whichsupplies the rate of change of localizer signal to the cross-coursevelocity sensor 414 in place of the heading signal used in the VOR mode.The cross-course velocity sensor is for the purpose of preventingpremature engagement in the approach mode control parameters. Therefore,unless the rate of approach is below said predetermined minimum theapproach mode cannot be engaged. In order that the third condition besatisfied, turn function selector knob 19 is rotated to its glide pathposition, thereby arming the coupler for automatic glide path beamcoupling. This arming is indicated to the pilot by the glide path armlight 416. It will be noted that should the turn function selector knob19 be inadvertently moved to the glide path position when no ILSfunction has in fact been selected, nothing will happen inasmuch as ILSswitch K-1453 open-circuits the glide slope interlock circuits.

Assuming now that the craft is within the 50 fia. portion of the beamand is being maintained on the beam by the radio and radio rate signals,and that the craft is approaching the glide slope beam from theunderside thereof. Vertical beam sensor 414 (this may be the samecircuit employed for the over-the-station sensor) will prevent prematureengagement of the glide slope control and also premature switching ofthe lateral chanel to its approach parameters until the aircraft isabout to intercept the center of the glide slope beam. The vertical beamsensor derives its input signal from the glide path receiver 311 throughnow closed relay switch K-143-4. Its output, when the glide slope beamerror signal drops below a predetermined value, for example, millivolts,controls vertical beam sensor relay Isl-146 which, in turn, energizesglide path engage relay K-147. The latter relay fulfills the thirdcondition mentioned above by closing switch K-M--Z in the approach relayK-l42 energization circuit. As approach relay K-MZ is energized, aholding circuit, through ILS relay switch K-14S-3 and approach relayswitch M2-l, is established so that further operation of thecross-course velocity detector switch K-ldd-l will not deenergize theapproach relay. It will also be noted that a second approach relay K-lZZis energized through approach relay switch K-142-2.

The craft is now in its final approach maneuver and in order to providevery tight beam coupling various gains in the approach coupler arechanged. In the first place, the gain of the displacement follow-up loop4% is changed by operation of approach relay switch K-lZZ-l in adirection to increase the response of the follow-up loop to lateral beamdisplacements. At the same time, the magnitude of the displacementsignal supplied to the autopilot through amplifier 413 is increasedthrough approach relay switch K-lZZ-S, and finally, the magnitude of therate signal supp-lied to the autopilot is also increased throughoperation of approach relay switch {(-122-2. -In this manner, the craftis controlled to precisely follow the localizer beam during the approachmode. Since large bank angles are to be avoided in the final approachconfiguration, the bank limiter is adjusted by means of switch K-122-4to decrease the limits imposed upon the sum of the radio and radio rat-esignals, that is, to reduce the magnitude of the sum signal which it canpass. :It will be further noted that any standoff errors |which mayaccumulate during the approach are eliminated through the operation ofthe roll integrator 11 which is rendered responsive to any standoffdisplacement error through approach relay switch K-l42-3.

-In `order to increase the yaw stability of the craft during the aproachmaneuver, a signal from the directional gyro 167 (FIG. 1B) is insertedinto the rudder channel in such a manner that short term yaw gusts areopposed while long term t-urn commands by the radio are unopposed. Thisis accomplished through yaw follow-up high tachometer gradient relayK-l7l (FIG. 14) which is energized simultaneously with the energizationof approach relay K-122. Energization of this relay serves throughswitch K-17f1-1 to decrease the gain through the yaw follow-up loop 108so that it can readily follow up on long term gyro signals, yet at thesame time be unable to follow up on short term gyro signal. Thus, anyshort term deviations of the craft in yaw, such as produced by lateralgusts, etc., are fed directly to the rudder servo system throughapproach relay switch K-142-4.

During the approach mode, the elevator channel of the autopilot iscontrolled in accordance with displacement of the craft vertically fromthe center of the glide slope beam. Any deviation of the craft from thebeam is detected by glide slope receiver 3H which supplies an errorsignal to the elevator servo system through amplifier limiter 321 whichoperates in the glide slope mode, exactly as in the altitude controlmode, to provide a displacement reference command signal for theelevator servo system. Additionally, and again as in the altitudecontrol mode,

any persistent glide slope displacement signal is integrated out throughthe operation of the pitch command computer 302 which has been placed inits integrator mode of operation. The latter is accomplished throughvertical path relay K-lllZ becoming energized when glide path engagerelay K-147 became energized, i.e., through switch K-147-4 (FIG. 9);this, in turn, energizing pitch computer gear shift solenoid `MG--lflLInertial path damping is also provided in the approach mode as in thealtitude cont-rol mode through the operation of vertical path relayK-lllZ through its switch K-112-3, which supplies a signal in accordancewith the vertical acceleration of the craft, the latter signal beingsupplied to the pitch command computer 302 where it is integrated andappears in its output as a rate of change of height signal. Therefore,the craft is controlled to precisely follow the glide slope beam throughthe combined effect o-f the displacement of the craft on the beam, therate of change of height of the craft, and the time integral of thedisplacement.

It will be noted that in the radio coupling modes of operation of theautopilot, the switching sequences which occur automatically frominitial bracketing to on course or approach are dependent solely uponthe position and/or movement of the craft with respect to the beam,thereby providing a very positive coupling to the beam.

From the foregoing specification, it will be clear that the autopilotdisclosed herein is similar in a great many respects to that disclosedin the above-identified copending parent application and yet there areherein disclosed many novel features not specifically set forth in theparent application. However, although specific and detailed disclosuresof the autopilot of `the present invention have been set forth, it isclearly apparent that many changes could be made in this specificconstruction and many widely different embodiments could be constructedwithout departing from the true scope and spirit thereof. Therefore, itis intended tthat all matter contained in the above description or shownin the accompanying drawings should be interpreted as illustrative andnot in a limiting sense.

What is claimed is:

l. In an automatic pilot for aircraft adapted to be controlled so as toapproach a radio beam and thereafter to maintain said beam, thecombination comprising means for supplying a plurality of signals eachproportional to the displacement of the craft fro-m said radio beam andeach having a different predetermined value for the same displacement,means for supplying a signal proportional to the heading of the craftwith respect to the bearing of said beam, first sensor means responsiveto a first magnitude of one of said radio displacement signals forcombining another of said radio displacement signals and said headingsignal, and second sensor means responsive to a second magnitude of said`one radio displacement signal for combining a further radiodisplacement signal and said heading signal whereby the sensitivity ofthe control of said craft by said radio and heading signals is varied inaccordance with the displacement thereof from the eam.

2, In an automate pilot for aircraft aadpted to be controlled so as toapproach a radio beam and thereafter to maintain said beam, thecombination comprising means for supplying a plurality of signals ea-chproportional to the displacement of the craft from said radio beam andeach having a different predetermined value for the same displacement,means for supplying a signal in accordance with the rate of approach ofthe craft with respect to the bearing of said beam, first sensor meansresponsive to a first magnitude of one of said radio displacementsignals for combining another of said radio displacement signals andsaid rate of approach signal, and second sensor means responsive to asecond magnitude of said one radio displacement signal for combining afurther radio displacement signal and said rate of approach signalwhereby the 23 'sensitivity of the control of said craft is varied as afunction of the displacement of the craft from the beam.

3. Apparatus as set forth in claim 2, wherein the means for providingsaid one displacement signal comprises a shaft positioned in accordancewith the displacement of the craft from the radio beam, wherein saidfirst sensor means comprises a first sector switch positioned by saidshaft having an elongated sector representative of a relatively largemagnitude of said beam displacement, and wherein said second `sensormeans comprises a second sector switch having a relatively short sectorrepresentative of a relatively small displacement of the craft from saidbeam.

4. 'In ran automatic pilot for aircraft adapted to be controlled so asto approach a radio-defined beam and thereafter to maintain said beamand wherein the relative values of the control signals controlling saidautopilot are changed upon substantial completion of said approach, thecombination comprising means for providing a signal proportional todisplacement of said craft from said beam, means for supplying a signalproportional to the heading of the aircraft with respect to the bearingof said ibeam, means for controlling the aircraft in accordance with thealgebraic sum of said signals whereby to cause said craft toasymptotically approach said beam, means responsive to a predeterminedlow value of said heading signal for supplying an output proportional tosaid value, and means responsive to said output signal for changing therelative values of the signals controlling said autopilot.

5. In yan automatic pilot for aircraft adapted to be controlled so as toapproach a radio defined beam and thereafter to maintain said beam andwherein the relative values of the control signals controlling saidautopilot are changed upon completion of said approach, the combinationcomprising means for providing a signal proportional to displacement ofsaid craft from said beam, means for supplying a signal proportional `tothe rate of approach of the aircraft with respect to said beam, meansfor controlling the 4aircraft in accordance with the algebraio sum ofsaid signals whereby to cause said craft to asymptotically approach saidbeam, means responsive to a predetermined low value of said rate signalfor supplying 'an output proportional to said value, and meansresponsive to said output signal for changing the relative values of thesignals controlling said autopilot.

Haskins Apr. 5, 1960 Carpenter July 25, 1961

1. IN AN AUTOMATIC PILOT FOR AIRCRAFT ADAPTED TO BE CONTROLLED SO AS TOAPPROACH A RADIO BEAM AND THEREAFTER TO MAINTAIN SAID BEAM, THECOMBINATION COMPRISING MEANS FOR SUPPLYING A PLURALITY OF SIGNALS EACHPROPORTIONAL TO THE DISPLACEMENT OF THE CRAFT FROM SAID RADIO BEAM ANDEACH HAVING A DIFFERENT PREDETERMINED VALUE FOR THE SAME DISPLACEMENT,MEANS FOR SUPPLYING A SIGNAL PROPORTIONAL TO THE HEADING OF THE CRAFTWITH RESPECT TO THE BEARING OF SAID BEAM, FIRST SENSOR MEANS RESPONSIVETO A FIRST MAGNITUDE OF ONE OF SAID RADIO DISPLACEMENT SIGNALS FORCOMBINING ANOTHER OF SAID RADIO DISPLACEMENT SIGNALS AND SAID HEADINGSIGNAL, AND SECOND SENSOR MEANS RESPONSIVE TO A SECOND MAGNITUDE OF SAIDONE RADIO DISPLACEMENT SIGNAL FOR COMBINING A FURTHER RADIO DISPLACEMENTSIGNAL AND SAID HEADING SIGNAL WHEREBY THE SENSITIVITY OF THE CONTROL OFSAID CRAFT BY SAID RADIO AND HEADING SIGNALS IS VARIED IN ACCORDANCEWITH THE DISPLACEMENT THEREOF FROM THE BEAM.
 4. IN AN AUTOMATIC PILOTFOR AIRCRAFT ADAPTED TO BE CONTROLLED SO AS TO APPROACH A RADIO-DEFINEDBEAM AND THEREAFTER TO MAINTAIN SAID BEAM AND WHEREIN THE RELATIVEVALUES OF THE CONTROL SIGNALS CONTROLLING SAID AUTOPILOT ARE CHANGEDUPON SUBSTANTIAL COMPLETION OF SAID APPROACH, THE COMBINATION COMPRISINGMEANS FOR PROVIDING A SIGNAL PROPORTIONAL TO DISPLACEMENT OF SAID CRAFTFROM SAID BEAM, MEANS FOR SUPPLYING A SIGNAL PROPORTIONAL TO THE HEADINGOF THE AIRCRAFT WITH RESPECT TO THE BEARING OF SAID BEAM, MEANS FORCONTROLLING THE AIRCRAFT IN ACCORDANCE WITH THE ALGEBRAIC SUM OF SAIDSIGNALS WHEREBY TO CAUSE SAID CRAFT TO ASYMPTOTICALLY APPROACH SAIDBEAM, MEANS RESPONSIVE TO A PREDETERMINED LOW VALUE OF SAID HEADINGSIGNAL FOR SUPPLYING AN OUTPUT PROPORTIONAL TO SAID VALUE, AND MEANSRESPONSIVE TO SAID OUTPUT SIGNAL FOR CHANGING THE RELATIVE VALUES OF THESIGNALS CONTROLLING SAID AUTOPILOT.